The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
An aircraft is propelled by several turbojet engines each housed within a nacelle, each nacelle further accommodating an assembly of additional actuating devices linked to its operation and ensuring various functions when the turbojet engine is in operation or stopped.
The modern nacelles are intended to accommodate a bypass turbojet engine capable of generating, by means of the fan blades in rotation, a flow of hot gases (also called primary flow) and a flow of cold air (also called secondary flow) witch circulates outside the turbojet engine through an annular passage, also called flow path, formed between two concentric walls of the nacelle. The primary and secondary flows are ejected from the turbojet engine by the rear of the nacelle.
A turbojet engine nacelle generally has a tubular structure including, from upstream to the downstream (relative to the direction of the cold and hot flows):                a front section, or air inlet, located in front of the turbojet engine;        a median section, intended to surround a fan module of the turbojet engine;        a rear section, intended to surround a high-pressure module, which in particular includes the combustion chamber of the turbojet engine, and generally embarking thrust reversal means;        an ejection nozzle, whose outlet is located downstream of the turbojet engine.        
The rear section generally has a fixed external structure, called “Outer Fixed Structure” (OFS), which defines, with a concentric internal structure, called “Inner Fixed Structure” (IFS), a downstream portion of the secondary flow path serving to channel the flow of cold air. The rear section is positioned downstream of a fan module of the turbojet engine which comprises in particular: a fan casing (inside which the fan is contained) and an intermediate casing. The intermediate casing includes a hub and an outer annular casing, as well as radial link arms therebetween.
Each propulsion unit of the aircraft is thus formed by a nacelle and a turbojet engine, and is suspended from a fixed structure of the aircraft, for example under a wing or on the fuselage, by means of a pylon or a mast fastened to the turbojet engine or to the nacelle.
It is thus observed that an aircraft propulsion unit integrates functional subassemblies likely to enter in relative movements, and between which it is suitable to manage the sealing.
In particular, it is important that the rear section of the nacelle, which delimits the secondary flow path, can be correctly aligned with the intermediate casing, with which it cooperates to channel the flow of cold air without leakage and without aerodynamic losses. Such a leakage would be particularly harmful, because a nacelle is designed and dimensioned to withstand the pressure exerted by the cold flow, in the case where it is correctly channeled. In contrast, the nacelle is not designed to withstand the forces generated by the pressure exerted by an air leakage of the secondary flow path towards the turbojet engine. Such a leakage can thus lead to a detachment of the inner structure of the nacelle. In view of these constraints, it is therefore essential to provide for a sealing barrier between the upstream portion of the rear section and the turbojet engine, in order to prevent any leakage of the secondary flow path towards the turbojet engine.
However, the sealing between the two covers and the turbojet engine presents a particular problem. First of all, the elements constituting the rear section of the nacelle are, in operation, animated by axial and radial movements relative to the turbojet engine. Given the large dimension of the concerned parts, these relative movements can, in operation, result in important displacements.
On the other hand, in operation, during the flight phases, the engine also undergoes deformations. In particular, the torsional forces generated by the rotation at very high speed of the fan blades lead the engine to be deformed about its longitudinal axis. This torsional movement, known under the name of “fan twist,” leads to an angular offset between the front part (the fan module, including in particular the intermediate casing) and the rear part (including in particular the combustion chamber) of the engine.
This angular offset is consequently also induced between the intermediate casing and the inner fixed structure. A gasket interposed between the inner fixed structure and the turbojet engine must therefore create a sealing barrier whatever the relative position of the inner fixed structure with respect to the turbojet engine, and for that, it must have a high crushing amplitude.
However, even by providing for such a gasket, the angular deformation of the engine in operation has severe disadvantages, among which is the reduction of the aerodynamic qualities of the secondary flow path. Indeed, the alignment of the inner fixed structure of the intermediate casing, which is correct when the engine is stopped, may become defective in flight. Indeed, the angular offset (about the longitudinal axis of the engine) between the inner fixed structure and the engine results in a deviation between some engine walls located in the flow of cold air when the engine is in operation, and which should normally be aligned with corresponding walls of the inner fixed structure. These walls are for example constituted by the outer surfaces of the link arms of the intermediate casing (and in particular those located in the positions called “6h00” and “12h00” positions. These alignment deviations generate a discontinuity of the aerodynamic lines of the secondary flow path, which greatly reduces the aerodynamic qualities of the secondary flow path.